Gas turbine engines, such as those which power aircraft and industrial equipment, employ a compressor to compress air that is drawn into the engine and a turbine to capture energy associated with the combustion of a fuel-air mixture.
There are frequently several flanges located at or near the exterior of the engine that separate the various sections of the engine. For example, and referring to the system 200 of FIG. 2, a flange 204 serves to separate a diffuser case 210 and a high pressure turbine (HPT) case 216. The flange 204 includes one or more holes to seat one or more fasteners (e.g., a bolt 222 and a nut 228) to couple the cases 210 and 216 to one another.
The portion of the engine in proximity to the flange 204 is typically one of the hottest, as the portion is located radially outboard of a combustion chamber 234. The flange 204 features two distinct areas where the radial interference of two parts form an interference fit; this occurs at the fully circumferential landing between the diffuser case 210 and the HPT case 216. The radially inner surface of this landing also provides a mating face to a first stage HPT turbine vane support 240 of a first stage HPT vane 246.
The arrangement of the system 200 results in the radially inner portion of the flange 204 being at a much higher temperature than the radially outer portion of the flange 204 where the holes are that seat the bolts 222. The temperature gradient could vary as much as, e.g., 400° Fahrenheit depending on the power settings of the engine. This temperature gradient results in thermally driven stress in the flange 204, which results in a low lifetime (frequently referred to in the art as a low cycle fatigue (LCF)) limit in the diffuser case 210.